The present invention relates generally to thermal insulation, and, more particularly, to a structure for providing thermal insulation, structural load accommodation, and noise attenuation for aircraft, missiles, and other aerospace systems having surfaces exposed to high temperatures.
In ducting hot exhaust gases from the engines of high-speed aircraft, it is necessary to protect the surrounding structure from excessively high temperatures. The required protection is typically accomplished through the use of multilayered structures that have a refractory material on the surface exposed to the high-temperature gases. Passageways for the circulation of cooling fluid are often provided between the refractory insulation layer and the underlying substructure of the aircraft. In such arrangements, a refractory metal, such as columbium, may be used to fabricate the layer that encounters the exhaust gases. A substructure consisting of one or more spaced-apart layers of superalloy underlie the refractory metal. While these arrangements are effective, they are not entirely satisfactory from the standpoint of cost and efficiency. To provide and circulate the coolant, it is necessary to use a portion of the engine output. To meet the cooling requirements of a high-powered engine exhaust duct, a significant energy expenditure is required. Commonly, as much as 50 percent of the available output may be needed for cooling purposes. This energy expenditure dramatically reduces the overall efficiency of a high-shaft-power or thrust propulsion system. A further disadvantage of these arrangements is the added weight and expense attributable to the materials used for the insulating layer and the underlying substructure.
The outer surfaces of space vehicles are also subjected to high temperatures during reentry. To protect these vehicles from the heat generated during reentry, various arrangements utilizing ablative materials have been heretofore proposed, often in combination with transpirational cooling systems. Because of their ablative characteristics, such systems are not well suited for insulating the surfaces of reusable spacecraft. Accordingly, efforts have been directed toward the development of nonablative insulating structures. In the present space shuttle, for example, portions of the shuttle exterior surface are insulated with a plurality of ceramic tiles that are arranged in a closely spaced, ordered array. To provide the required fit, each tile must be precision cut from a carefully formed fused ceramic blank. To form the blanks, silica fibers and other ceramic components are initially mixed into a slurry and cast into blocks. After drying, the blocks are sintered at high temperatures to form strong ceramic bonds between the overlapping fibers. The blocks are sawn thereafter into the smaller blanks that are subsequently configured into the final tiles by a numerically controlled mill. Once prepared each tile is individually secured in place via a manual procedure. This involves bonding the tiles to a felt strain isolation pad with a high-temperature adhesive, then adhesively bonding the pad to the underlying metallic substructure.
During takeoff and reentry, nonuniform temperature gradients exist across the surface of the space shuttle, which is insulated with these ceramic tiles. The fused ceramic structure of the tiles is poorly resistant to shear forces, and, thus, poorly resistant to the forces occasioned by the differential surface temperature distribution. Accordingly, to prevent breakage of the ceramic insulation, the tiles must be limited to small sizes, generally less than ten inches on a side. While this isolates the loads applied to the insulation, it does so by exacting an extremely high cost in terms of parts production and assembly onto the spacecraft.
For applications other than the space shuttle, these high costs may render the ceramic tile approach unfeasible. Where the application dictates the use of an unbroken, i.e., continuous, insulation surface, the ceramic tile approach would be entirely unworkable. Since the cast ceramic blocks shrink substantially and in an irregular manner during the sintering operation, the use of such fused fibrous ceramic materials is also not well suited for forming insulating surfaces that have a substantial degree of curvature.
The present invention provides an arrangement that overcomes the disadvantages of the developments described above. In particular, the invention provides an insulating structure that not only affords protection against high temperature, but also accommodates loads. This load-bearing capability enables a reduction in the amount of substructure required to support the insulation, and, thus, contributes to an overall reduction in the weight of the craft. The invention provides, as well, a thermal insulating structure that also functions as a noise attenuator and, thus, is particularly well suited for use in an aircraft engine exhaust duct. An important aspect of the invention is the provision of a thermal insulating structure that is fabricated by casting a fibrous ceramic insulation material onto a honeycomb core. In accordance with a particular aspect of the invention, this manufacture of the insulating structure is enabled through the use of a honeycomb core having a perforated buried septum. As a result of the support provided by the honeycomb core to the ceramic insulation, structures having continuous insulating surfaces may be fabricated in a wide variety of shapes and configurations